Gas turbine engines, particularly those used in aircraft applications, place an unusual demand on materials used in the high temperature portions of the engine. For example, in the turbine section of an engine, a 1100.degree.-1500.degree. C. gas stream impinges on a multiplicity of air cooled blades mounted around the rim of a spinning disk. The parts are thus subjected to a severe combination of high axial loads and high temperatures. Over the past 40 years there has been considerable technical effort directed toward producing improved materials and designs to optimize the performance of the high temperature components in a gas turbine engine. Initially, forged stainless steels and nickel chromium alloys were used; later, these were superseded by gamma prime containing superalloys. However, as the gamma prime content of alloys was increased over time, the alloys became non-forgeable; this led to the use in the 1960's of cast turbine blades. In the 1970's, advances in casting technology led to the introduction of directionally solidified columnar grain alloys having greatly improved creep strength, as disclosed in U.S. Pat. No. 3,260,505 to VerSnyder. Currently, directionally solidified single crystal alloy parts, as disclosed in U.S. Pat. No. 3,494,709 to Piearcey, are in commercial use.
The foregoing progression of materials and processes has enabled gas turbine blades to be operated at progressively higher material temperatures and stresses. Still, there is demand for further improvements and the present invention seeks to satisfy these needs. Considering in more detail a gas turbine blade, it will be found that the airfoil portion of a blade has the highest material temperatures, resulting in a tendency toward failure by creep or thermal fatigue. The root portion of the blade, where it is held in the rim of the necessarily cooler disk, tends to be at a lower temperature. The root region tends toward failure due to mechanical fatigue or tensile yielding.
Nonetheless, if improved strength can be provided in a material for any particular hot or cold regime, then the designer of the gas turbine blade can reduce the amount of material which is needed in the portion subjected to that particular regime. When a part is thereby made lighter, there may be a magnified benefit in reduction in weight of all the structure supporting the turbine blade. Alternatively the life of a part may be extended while maintaining it at its original configuration.
The metallurgical design of a single crystal alloy, as with most superalloys, is necessarily a compromise. Therefore, while there have been important advances in the chemistry and metallurgy of single crystals to produce improved low and high temperature properties, still further opportunities for improving performance are of economic interest.
Previously, Owczarski et al disclosed in U.S. Pat. No. 3,642,543 that the yield strength of certain single crystal alloys could be improved by hot rolling the alloys to about 40% reduction. At least 15% reduction was said to be necessary to get any improvement. However, the dimensions of gas turbine blades must be very precise and they are usually cast and machined to quite high tolerances. The intricate internal passages that characterize modern air cooled turbine blades usually must be cast into place for economy; machining such passages is generally infeasible and often impossible. Therefore, the previously described rolling procedure is not really pertinent to the present requirement.